The present invention relates generally to gas turbine engines, and, more specifically, to compressors or fans therein.
In a turbofan aircraft gas turbine engine, air is pressurized in a fan and compressor during operation. The fan air is used for propelling an aircraft in flight. The air channeled through the compressor is mixed with fuel in a combustor and ignited for generated hot combustion gases which flow through turbine stages that extract energy therefrom for powering the fan and compressor.
A typical turbofan engine includes a multistage axial flow compressor which pressurizes the air sequentially to produce high pressure air for combustion. The compressed air is diffused and decelerates as it is compressed. Compressor airfoils must therefore be designed to reduce undesirable flow separation which would adversely affect stall margin and efficiency.
Conversely, combustion gases are accelerated through the turbine stages, and the turbine blades have different aerodynamic designs for maximizing efficiency of energy extraction.
Fundamental in compressor design is efficiency in compressing the air with sufficient stall margin over the entire flight envelope of operation from takeoff, cruise, and landing.
However, compressor efficiency and stall margin are normally inversely related with increasing efficiency typically corresponding with decrease in stall margin. The conflicting requirements of stall margin and efficiency are particularly demanding in high performance military engine applications which require high level of stall margin typically at the expense of compressor efficiency, as opposed to less demanding commercial applications.
Maximizing efficiency of compressor airfoils is primarily effected by optimizing the velocity distributions over the pressure and suction sides of the airfoil. However, efficiency is typically limited in conventional compressor design by the requirement for a suitable stall margin. Any further increase in efficiency typically results in a reduction in stall margin, and, conversely, further increase in stall margin results in decrease in efficiency.
High efficiency is typically obtained by minimizing the wetted surface area of the airfoils for a given stage to correspondingly reduce airfoil drag. This is typically achieved by reducing airfoil solidity or the density of airfoils around the circumference of a rotor disk, or by increasing airfoil aspect ratio of the span to chord lengths.
For a given rotor speed, this increase in efficiency reduces stall margin. To achieve high levels of stall margin, a higher than optimum level of solidity and/or lower than optimum aspect ratios may be used, along with designing the airfoils at below optimum incidence angles. This reduces axial flow compressor efficiency.
Increased stall margin may also be obtained by increasing rotor speed, but this in turn reduces efficiency by increasing the airflow Mach numbers, which increases airfoil drag.
And, compressor blades are subject to centrifugal stress which is affected by aerodynamic design. Peak stress must be limited for obtaining useful blade life, and this in turn limits the ability to optimize aerodynamic performance.
Accordingly, typical compressor designs necessarily include a compromise between efficiency and stall margin favoring one over the other, which are further affected by allowable centrifugal stress.
It is, therefore, desired to further improve both compressor efficiency and stall margin while limiting centrifugal stress for improving gas turbine engine compressor performance.